The embodiments described herein relate generally to gas turbine engines and more particularly relate to an axial compressor endwall treatment for a gas turbine engine and a method for controlling leakage flow and circumferential flow non-uniformities therein.
As is known, an axial compressor for a gas turbine engine may include a number of stages arranged along an axis of the compressor. Each stage may include a rotor disk and a number of compressor blades, also referred to herein as rotor blades, arranged about a circumference of the rotor disk. In addition, each stage may further include a number of stator blades, disposed adjacent the rotor blades and arranged about a circumference of the compressor casing.
During operation of a gas turbine engine using a multi-stage axial compressor, a turbine rotor is turned at high speeds by a turbine so that air is continuously induced into the compressor. The air is accelerated by the rotating compressor blades and swept rearwards onto the adjacent rows of stator blades. Each rotor blade/stator blade stage increases the pressure of the air. In addition, during operation a portion of the compressed air may pass downstream about a tip of each of the compressor blades and/or stator blades as a leakage flow. Such stage-to-stage leakage of compressed air as leakage flow may affect the stall point of the compressor.
Compressor stalls may reduce the compressor pressure ratio and reduce the airflow delivered to a combustor, thereby adversely affecting the efficiency of the gas turbine. A rotating stall in an axial-type compressor typically occurs at a desired peak performance operating point of the compressor. Following rotating stall, the compressor may transition into a surge condition or a deep stall condition that may result in a loss of efficiency and, if allowed to be prolonged, may lead to failure of the gas turbine.
The operating range of an axial compressor is generally limited due to weak flow in rotor tips, where the specific rotor stall point is determined by the operating conditions, circumferential flow non-uniformities and compressor design. Prior attempts to increase the range of this operation and increase the stall margin have included flow control based techniques such as plasma actuation and suction/blowing near a blade tip. However, such attempts significantly increase compressor complexity and weight. Other attempts include end-wall treatments such as circumferential grooves, axial grooves, or the like. These end-wall treatments do not rotate with the rotor, and have a fixed relative position (both axially and circumferentially) to the upstream stationary blade-row. In addition, known end-wall treatments are predominantly oriented in the axial direction, and are all geometrically identical circumferentially about the entire annulus. It is known that the presence of upstream blades or struts introduce the circumferential flow non-uniformities. As such, these geometrically identical end-wall treatments are not designed to exploit/leverage the circumferentially non-uniform flows introduced by the upstream blade-row and are not an optimal arrangement to improve stall margins.
Thus, there is a desire for an improved axial compressor for a gas turbine engine and a method for controlling leakage flow about one or more blade tips in the presence of circumferential flow non-uniformities. Specifically, such a compressor may control leakage of compressed air through a carefully designed endwall treatment proximate the rotor and/or stator blades that provides desired recirculation of the leakage flow and addresses the circumferential flow non-uniformities. Control of such leakage and circumferential flow non-uniformities may increase operating range and stall margin of the compressor and the overall gas turbine engine while minimizing the detrimental impact on design point efficiency.